1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to large turbine airfoils with a cooling circuit.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine is very efficient machine that converts the chemical energy of a burning fuel into mechanical energy. An industrial gas turbine (IGT) engine is used in power plants to drive an electric generator to produce electric power. The efficiency of a gas turbine engine can be increased by increasing the high temperature gas flow that enters the turbine. It is a very important design feature to provide for the first stage stator vanes and rotor blades to have as high of a high heat resistance as possible by using high temperature resistant materials in combination with internal and film cooling of the airfoils (vanes and blades).
In the recent history of industrial gas turbine engines, because the turbine inlet temperature was not too high, only the first and second stages of stator vanes and rotor blades required cooling. With the recent improvement in airfoil materials and cooling, the turbine inlet temperature has increased to the point where the third stage and even the fourth stage airfoils require cooling in order to have a long life time. Even though the gas flow temperature acting on the fourth stage rotor blades is not high enough to melt the blades, the temperature is high enough to result in creep and other thermal effects on the blades that will shorten the blade's life time in operation. It is desirable to design a fourth stage blade for 96,000 hours of operation in order to reduce the high cost of replacing these blades.
Also, because the inlet temperature to the fourth stage rotor blades is high enough, the size of these blades must be increased in order to maximize the energy extracted from the hot gas flow. As the fourth stage rotor blades increase in length, the twisting that results on the airfoil causes problems with creating the cooling holes within the blades. Straight or radial holes cannot be drilled from tip to root because of the twist. In addition, the core ties used in casting the internal passages within these blades are easily damaged in the molds, and as a result defective blades are cast.
Another problem with large turbine rotor blades is the effect of such a relatively large mass due to rotation of the blade under extreme high temperature. The centrifugal force on the rotating blade in addition to the high temperature will lead to creep problems or to the blade untwisting due to deformation. The aero-performance of the blade and well as the remaining life of the blade will both decrease.
Prior art cooling of large turbine rotor blade is achieved by drilling radial holes into the blade from the blade tip and root sections. Limitations of drilling a long radial hole from both ends of the airfoil increases for a large and highly twisted and tapered blade. Reduction of available airfoil cross sectional area for drilling radial holes is a function of the blade twist and taper. Higher airfoil twist and taper yield a lower available cross sectional area for drilling radial cooling holes. Cooling of the large, highly twisted and tapered blade by this manufacturing technique will not achieve the optimum blade cooling effectiveness. Especially lacking is cooling for the blade leading edge and trailing edge. This prevents the use of such blades in a high firing temperature application as well as a low cooling flow design. FIG. 1 shows a prior art turbine airfoil for a large rotor blade with a cooling flow design that uses the drilling of radial cooling holes, which is U.S. Pat. No. 6,910,864 issued to Tomberg on Jun. 28, 2005 and entitled TURBINE BUCKET AIRFOIL COOLING HOLE LOCATION, STYLE AND CONFIGURATION.
U.S. Pat. No. 5,993,156 issued to Bailly et al on Nov. 30, 1999 and entitled TURBINE VANE COOLING SYSTEM discloses a turbine vane cooling system in which the vane includes a cooling air supply orifice (23 in this patent in FIGS. 2 and 12) located in the vane root which then splits up into two flows (B1 and B2 in FIG. 12 of this patent), with one path along the pressure side and the other path along the suction side. The two paths then are combined into a central cavity (# 13 and 15 in FIG. 12 of this patent), and then passes through an aperture (# 18 in this patent) located at the base and into a trailing edge channel (# 16 in this patent) in which cooling air outlet slots (# 19 in this patent) discharge the cooling air out from the vane.
U.S. Pat. No. 5,779,447 issued to Tomite et al on Jul. 14, 1998 and entitled TURBINE ROTOR discloses a rotor blade with a cooling circuit having a lower cavity (# 4 in this patent) with pin fins extending across the cooling passages formed by ribs (# 14 in this patent), and an upper portion of the blade having a plurality of holes (# 15 in this patent) extending from the lower cavity to the blade tip.
U.S. Pat. No. 6,152,695 issued to Fukue et al on Nov. 28, 2000 and entitled GAS TURBINE MOVING BLADE discloses a rotor blade in FIG. 1 of this patent with an inner cavity (# 10 in this patent) separated by ribs with pin fins extending across the cavity that extends from the root to the blade tip, and another embodiment in FIG. 15 of this patent in which the cavity stops short of the blade tip in which radial holes continue until the blade tip.
It is therefore an object of the present invention to provide for a large high tapered turbine rotor blade with internal air cooling circuit that will provide adequate cooling of the blade while also being easily cast without significant errors in the casting.
It is also another object of the present invention to provide for a large turbine rotor blade with an increase in the AN2 of the prior art blades.